The efficiency of gas turbine engine may be improved by raising the turbine gas inlet temperature. In the present state of the art, this temperature is limited because both the blades and the rotor disk of the turbine rotor are metallic and cannot withstand the gas temperatures above certain maximum values. Ceramic materials are currently under investigation for use in making turbine blades. One of the properties of ceramic material is its very low heat transfer. This results in high thermal gradients within the parts of the turbine rotor when ceramic blades are used. A full sized blade, i.e., one having a chord of about 4.0 inches, in the first stage of a typical gas turbine engine would crack due to bowing along the chord length at the root of the blade. The bowing would be caused by the temperature gradient from that area of the blade exposed to the hot gas flow to the blade area buried within the rotor disk or intermediate attachment piece connecting the blade to the disk.
One solution that has been proposed is to reduce the blade section by 50%, thereby reducing its chord accordingly while maintaining the original blade height. This results in a blade section which is much thinner relative to the length, a problem which is not desirable. The obvious disadvantage is that, with such a smaller blade section, the effect is to double the tolerances of the blade unless the tolerances are also reduced by 50%. This approach is, therefore, impractical.
In view of the foregoing problem, a need has arisen for an improved turbine blade to compensate for bowing in the base of the blade while maintaining a full sized blade section.